Fuel nozzle with angled main injection ports and radial main injection ports

ABSTRACT

The present disclosure is directed to a combustor system including a fuel nozzle comprising a main fuel injector defining an angled main injection port. Each angled main injection port defines an inlet end and an outlet end. The outlet end is oriented downstream relative to the inlet end at an angle relative to a centerline axis of the fuel nozzle. The angled main injection port permits egress of fuel from a main fuel circuit to the combustion chamber. The fuel nozzle further defines a radial main injection port. The outlet end is oriented along an axial direction approximately equal relative to the inlet end and radially outward thereof relative to the centerline axis. The radial main injection port is extended perpendicular relative to the centerline axis.

FEDERALLY SPONSORED RESEARCH

This invention was made with government support under contract numberFA8650-07-C-2802 awarded by the U.S. Department of Defense. Thegovernment may have certain rights in the invention.

FIELD

The present subject matter relates generally to gas turbine enginecombustor assemblies. More particularly, the present subject matterrelates to combustor assembly structures affecting combustion dynamics.

BACKGROUND

Lean-burn combustor assemblies for gas turbine engines (e.g., propulsionengines for aircraft) are generally susceptible to undesirablecombustion dynamics (e.g., acoustics, vibrations, and pressureoscillations resulting from heat release characteristics fromcombustion). Combustion dynamics may result in accelerated deteriorationand wear of combustor assemblies and gas turbine engines, thusincreasing maintenance and operational costs, reduced efficiencies andengine operability, and increase the risk of overall engine failure.

Thus, there is a need for a combustor assembly that mitigates undesiredcombustion dynamics.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

The present disclosure is directed to a combustor system including afuel nozzle comprising a main fuel injector defining an angled maininjection port. Each angled main injection port defines an inlet end andan outlet end. The outlet end is oriented downstream relative to theinlet end at an angle relative to a centerline axis of the fuel nozzle.The angled main injection port permits egress of fuel from a main fuelcircuit to the combustion chamber. The fuel nozzle further defines aradial main injection port. The outlet end is oriented along an axialdirection approximately equal relative to the inlet end and radiallyoutward thereof relative to the centerline axis. The radial maininjection port is extended perpendicular relative to the centerlineaxis.

In various embodiments, the fuel nozzle defines an annular fuel nozzleouter wall radially spaced apart from an annular outer boundary wallsurrounding a pilot fuel injector. The fuel nozzle outer wall defines anopening that permits a flow of air between the outer wall and the outerboundary wall. In one embodiment, the fuel nozzle defines an aperturethrough the outer wall, in which the aperture is defined in alignmentwith the angled main injection port and defines an angle with respect tothe centerline axis approximately equal to the angle of the maininjection port. In another embodiment, the fuel nozzle defines anaperture through the outer wall, in which the aperture is defined inalignment with the radial main injection port along a radial direction.

In still various embodiments, the fuel nozzle comprises a pilot fuelinjector. The pilot fuel injector comprises an annular inner walldefining a primary fuel orifice and an annular outer wall defining asecondary fuel orifice. A primary pilot supply line supplies fuelthrough the primary fuel orifice and a secondary pilot supply linesupplies fuel through the secondary fuel orifice. In one embodiment, thecombustor system further includes a fuel system comprising a pilotcontrol valve coupled to a pilot fuel conduit and operable to supply aflow of fuel thereto. The pilot fuel conduit is in fluid communicationwith the primary pilot supply line and the secondary pilot supply line.In another embodiment, the fuel system further includes a first mainvalve coupled to a first main fuel conduit and operable to supply a flowof fuel thereto. The first main fuel conduit is in fluid communicationwith a first main fuel circuit of the fuel nozzle.

In still various embodiments, the fuel nozzle defines a different volumeor cross sectional area of the angled main fuel injection port, theinlet end, the outlet end, or combinations thereof relative to theradial main fuel injection port. In one embodiment, the combustor systemdefines a plurality of fuel nozzles disposed in adjacent circumferentialarrangement, and wherein the plurality of fuel nozzles defines a firstfuel nozzle defining the angled main fuel injection port and a secondfuel nozzle defining the radial main fuel injection port. In anotherembodiment, the first fuel nozzle, the second fuel nozzle, and the fuelsystem together provide passive combustion dynamics attenuation defininga non-uniform flame structure and characteristic time given anapproximately equal flow of fuel to the main injection port from thefirst main fuel circuit.

In still yet various embodiments, the fuel system further includes afirst main valve coupled to a first main fuel conduit and operable tosupply a flow of fuel thereto, wherein the first main fuel conduit is influid communication with a first main fuel circuit of the fuel nozzle;and a second main valve coupled to a second main fuel conduit andoperable to supply a flow of fuel thereto independent of the first mainfuel conduit, wherein the second main fuel conduit is in fluidcommunication with a second main fuel circuit of the fuel nozzle. In oneembodiment, the combustor system defines a plurality of fuel nozzlesdisposed in adjacent circumferential arrangement, and wherein theplurality of fuel nozzles defines a first fuel nozzle defining theangled main fuel injection port and a second fuel nozzle defining theradial main fuel injection port. The first fuel nozzle, the second fuelnozzle, and the fuel system together provide active combustion dynamicsattenuation defining a non-uniform flame structure and characteristictime given a variable pressure, flow, or temperature of fuel to the maininjection port from the first main fuel circuit coupled to the firstfuel nozzle and the second main fuel circuit coupled to the second fuelnozzle.

In one embodiment of the combustor system, the fuel nozzle furtherincludes a pilot fuel injector in which each of the main fuel injectorand the pilot fuel injector are configured to receive a portion of afuel flow to each fuel nozzle.

Another aspect of the present disclosure is directed to a gas turbineengine including the combustor system. In various embodiments, thecombustor system includes a fuel nozzle comprising a main fuel injectordefining an angled main injection port, each angled main injection portdefining an inlet end and an outlet end. The outlet end of the angledmain injection port is oriented downstream relative to the inlet end atan angle relative to a centerline axis of the first fuel nozzle. Theangled main injection port permits egress of fuel from a main fuelcircuit to the combustion chamber. The fuel nozzle further defines aradial main injection port in which the outlet end is oriented along anaxial direction approximately equal relative to the inlet end andradially outward thereof relative to the centerline axis. The radialmain injection port is extended perpendicular relative to the centerlineaxis.

In various embodiments of the gas turbine engine, the engine furtherincludes a fuel system comprising a pilot control valve coupled to apilot fuel conduit and operable to supply a flow of fuel thereto. Thepilot fuel conduit is in fluid communication with the primary pilotsupply line and the secondary pilot supply line. In one embodiment, thefuel system further comprises a first main valve coupled to a first mainfuel conduit and operable to supply a flow of fuel thereto, wherein thefirst main fuel conduit is in fluid communication with a first main fuelcircuit of the fuel nozzle.

In still various embodiments of the gas turbine engine, the fuel nozzledefines a different volume or cross sectional area of the angled mainfuel injection port, the inlet end, the outlet end, or combinationsthereof relative to the radial main fuel injection port. In oneembodiment, the combustor system defines a plurality of fuel nozzlesdisposed in adjacent circumferential arrangement. The plurality of fuelnozzles defines a first fuel nozzle defining the angled main fuelinjection port and a second fuel nozzle defining the radial main fuelinjection port. The first fuel nozzle, the second fuel nozzle, and thefuel system together provide passive combustion dynamics attenuationdefining a non-uniform flame structure and characteristic time given anapproximately equal flow of fuel to the main injection port from thefirst main fuel circuit.

In still various embodiments of the gas turbine engine, a first mainvalve is coupled to a first main fuel conduit and operable to supply aflow of fuel thereto. The first main fuel conduit is in fluidcommunication with a first main fuel circuit. A second main valve iscoupled to a second main fuel conduit and operable to supply a flow offuel thereto independent of the first main fuel conduit. The second mainfuel conduit is in fluid communication with a second main fuel circuit.In one embodiment, the combustor system defines a plurality of fuelnozzles disposed in adjacent circumferential arrangement, and whereinthe plurality of fuel nozzles defines a first fuel nozzle defining theangled main fuel injection port and a second fuel nozzle defining theradial main fuel injection port, and wherein the first fuel nozzle, thesecond fuel nozzle, and the fuel system together provide activecombustion dynamics attenuation defining a non-uniform flame structureand characteristic time given a variable pressure, flow, or temperatureof fuel to the main injection port from the first main fuel circuitcoupled to the first fuel nozzle and the second main fuel circuitcoupled to the second fuel nozzle.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 provides a schematic cross-section view of an exemplary gasturbine engine according to various embodiments of the present subjectmatter.

FIG. 2 provides a schematic cross-section view of a combustor system ofthe gas turbine engine of FIG. 1, according to an exemplary embodimentof the present subject matter.

FIG. 3 provides a schematic cross-section view of a combustor system ofthe gas turbine engine of FIG. 1, according to another exemplaryembodiment of the present subject matter.

FIG. 4 provides a schematic cross-section view of a fuel nozzle assemblyof the combustor system of FIGS. 2-3, according to an exemplaryembodiment of the present subject matter.

FIGS. 5, 6, and 7 provide enlarged views of segments of the fuel nozzleassembly illustrated in FIG. 4.

FIG. 8 provides a schematic cross-section view of a second fuel nozzleassembly of the combustor system of FIGS. 2-3, according to an exemplaryembodiment of the present subject matter.

FIG. 9 provides a schematic cross-section view of a first fuel nozzleassembly of the combustor system of FIGS. 2-3, according to an exemplaryembodiment of the present subject matter.

FIG. 10 provides a schematic cross-section view of a portion of a mainfuel injector of a fuel nozzle assembly, according to an exemplaryembodiment of the present subject matter.

FIG. 11 provides schematic cross-section view of a portion of a mainfuel injector of a fuel nozzle assembly, according to another exemplaryembodiment of the present subject matter.

FIG. 12 provides an aft end view of a portion of a fuel nozzle outlet,according to an exemplary embodiment of the present subject matter.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first,” “second,” and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows and “downstream” refers to thedirection to which the fluid flows.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure, referred to herein as“engine 10.” As shown in FIG. 1, the engine 10 defines an axialdirection A (extending parallel to a longitudinal centerline 12 providedfor reference) and a radial direction R. In general, the turbofan 10includes a fan section 14 and a core turbine engine 16 disposeddownstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22. In other embodiments ofengine 10, additional spools may be provided such that engine 10 may bedescribed as a multi-spool engine.

For the depicted embodiment, fan section 14 includes a fan 38 having aplurality of fan blades 40 coupled to a disk 42 in a spaced apartmanner. As depicted, fan blades 40 extend outward from disk 42 generallyalong the radial direction R. The fan blades 40 and disk 42 are togetherrotatable about the longitudinal axis 12 by LP shaft 36. In someembodiments, a power gear box having a plurality of gears may beincluded for stepping down the rotational speed of the LP shaft 36 to amore efficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 1, disk 42 iscovered by rotatable front nacelle 48 aerodynamically contoured topromote an airflow through the plurality of fan blades 40. Additionally,the exemplary fan section 14 includes an annular fan casing or outernacelle 50 that circumferentially surrounds the fan 38 and/or at least aportion of the core turbine engine 16. It should be appreciated thatnacelle 50 may be configured to be supported relative to the coreturbine engine 16 by a plurality of circumferentially-spaced outletguide vanes 52. Moreover, a downstream section 54 of the nacelle 50 mayextend over an outer portion of the core turbine engine 16 so as todefine a bypass airflow passage 56 therebetween.

During operation of the engine 10, a volume of air 58 enters turbofan 10through an associated inlet 60 of the nacelle 50 and/or fan section 14.As the volume of air 58 passes across fan blades 40, a first portion ofthe air 58 as indicated by arrows 62 is directed or routed into thebypass airflow passage 56 and a second portion of the air 58 asindicated by arrows 64 is directed or routed into the LP compressor 22.The ratio between the first portion of air 62 and the second portion ofair 64 is commonly known as a bypass ratio. The pressure of the secondportion of air 64 is then increased as it is routed through the highpressure (HP) compressor 24 and into the combustion section 26, where itis mixed with fuel and burned to provide combustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

It will be appreciated that, although described with respect to turbofan10 having core turbine engine 16, the present subject matter may beapplicable to other types of turbomachinery. For example, the presentsubject matter may be suitable for use with or in turboprops,turboshafts, turbojets, industrial and marine gas turbine engines,and/or auxiliary power units.

FIG. 2 provides a schematic cross-sectional view of a combustor system100, e.g., for use in the gas turbine engine of FIG. 1, according to anexemplary embodiment of the present subject matter. As shown in FIG. 2,the combustor system 100 comprises a combustor 101 having a forward end101 a and an aft end 101 b. The combustor 101 further includes anannular inner liner 102 and an annular outer liner 104. The inner liner102 extends generally along the axial direction A between an upstreamend 106 and a downstream end 108. Similarly, the outer liner 104 extendsgenerally along the axial direction A between an upstream end 110 and adownstream end 112. Each of the inner liner 102 and the outer liner 104may be formed from a CMC material, as described in greater detail below,or from any other suitable material.

A combustor dome 114 extends generally along the radial direction Rbetween the upstream end 106 of the inner liner 102 and the upstream end110 of the outer liner 104. As shown in FIG. 2, the inner liner 102, theouter liner 104, and the combustor dome 114 define a combustion chamber116 therebetween. In some embodiments, the combustor dome 114 isintegral with the inner liner 102, i.e., the inner liner 102 and thecombustor dome 114 are integrally formed as a single piece structure,but in other embodiments, the combustor dome 114 is integral with theouter liner 104, i.e., the outer liner 104 and the combustor dome 114are integrally formed as a single piece structure. In still otherembodiments, the combustor dome 114 is formed separately from the innerliner 102 and the outer liner 104, or in yet other embodiments, thecombustor dome 114 is integral with both the inner and outer liners 102,104, e.g., at least a first portion of the combustor dome 114 may beintegral with the inner liner 102 and at least a second portion of thecombustor dome 114 may be integral with the outer liner 104. Thecombustor dome 114 may be formed from any suitable material, e.g., a CMCmaterial or a metallic material, such as a metal or metal alloy.

Further, the combustor system 100 includes a fuel nozzle assembly 117having a plurality of fuel nozzles 113 each defining a fuel nozzleoutlet 120 at an outlet end 119 of the fuel nozzle 113. In variousembodiments, the combustor system 100 includes a first fuel nozzle 118and a second fuel nozzle 115 each disposed in alternatingcircumferential arrangement around the longitudinal centerline 12. Aswill be discussed further below, the plurality of fuel nozzles 113generally provided may provide a flame non-uniformity that may mitigatecombustion dynamics. Furthermore, the flame non-uniformity may mitigatehigh cycle fatigue of the combustor system 100 or components thereof. Instill various embodiments, the fuel nozzle 113 generally provided mayreduce surface or internal temperatures of the inner liner 102, theouter liner 104, or both, thereby improving engine 10 durability.

Referring still to the combustor system 100, a main mixer 190 extendsabout the fuel nozzle outlet 120 as described in greater detail below.The fuel nozzle 113 is disposed through the combustor dome 114 such thatthe fuel nozzle outlet 120 is disposed at or adjacent the forward end101 a of the combustor 101 to direct a fuel-air mixture into thecombustion chamber 116. More particularly, the exemplary fuel nozzle 113is of a type configured to inject liquid or gaseous hydrocarbon fuelinto an airflow stream of the combustor system 100. The fuel nozzle 113is of a “staged” type, meaning it is operable to selectively inject fuelthrough two or more discrete stages, each stage being defined byindividual fuel flowpaths within the fuel nozzle 113.

The fuel flowrate may be variable within each of the stages. In theexemplary embodiment depicted in FIG. 2, the fuel nozzle 113 isconnected to a fuel system 122 that is operable to supply a flow ofliquid or gaseous fuel (or combinations thereof) at varying flowratesaccording to operational need. The fuel system 122 supplies fuel to apilot control valve 124 that is coupled to a pilot fuel conduit 126,which in turn supplies fuel to a primary pilot supply line 128 and asecondary pilot supply line 130 (FIG. 4, FIG. 8-9) within the fuelnozzle 113.

In one embodiment of the combustion system 100, such as generallyprovided in FIG. 2, the fuel system 122 supplies fuel to a first mainvalve 132 that is coupled to a first main fuel conduit 134, which inturn supplies a first main fuel circuit 136 (FIG. 4, FIG. 9) and asecond main fuel circuit 137 (FIG. 4, FIG. 8) of a first fuel nozzle 118and a second fuel nozzle 115, respectively, of the plurality of fuelnozzles 113.

In another embodiment of the combustion system 100, such as generallyprovided in FIG. 3, the fuel system 122 supplies fuel to a second mainvalve 133 that is coupled to a second main fuel conduit 135, which inturn supplies the second main fuel circuit 137 (FIG. 8) of the secondfuel nozzle 115. The first main valve 132 is coupled to the first mainfuel conduit 134, which in turn supplies the first main fuel circuit 136(FIG. 4) of the first fuel nozzle 118. As such, a flow, pressure, ortemperature of fuel to the first fuel nozzle 118 may be independentlyadjustable relative to the second fuel nozzle 115.

The independent adjustability of the flow, pressure, or temperature ofthe fuel to the first fuel nozzle 118 versus the second fuel nozzle 115,in combination with the varying geometries of a fuel injection port 184of each fuel nozzle 113 (e.g., FIGS. 10-11), enables fuel modulationbetween the two that may mitigate combustion dynamics. Furthermore, theindependent adjustability may reduce inner liner 102 or outer liner 104temperature, improve fuel/air mixing, reduce auto-ignition of the fuel,or combinations of these benefits. Additionally, independentadjustability may enable control or adjustability of combustion patternfactor (i.e., variations in temperature around the annulus of thecombustion chamber 116).

Referring now to FIG. 4, a cross-section view is provided of a portionof the fuel nozzle assembly 117. Additionally, FIGS. 5, 6, and 7 provideenlarged views of segments of a portion of the fuel nozzle 113illustrated in FIG. 4. For purposes of description, reference will bemade to a centerline axis CL of the fuel nozzle assembly 117. In someembodiments, the centerline axis CL is generally parallel to the axialcenterline 12 of the engine 10, but in other embodiments, the centerlineaxis CL may be at an angle relative to the engine axial centerline 12.The components of the illustrated fuel nozzle assembly 117 are disposedextending parallel to and surrounding the centerline axis CL, generallyas a series of concentric rings. For instance, a pilot fuel injector 138is disposed at or near the outlet 120 of the fuel nozzle 113 and isaligned with the centerline axis CL. As shown most clearly in FIG. 5,the pilot fuel injector 138 includes a generally annular inner wall 140that defines a primary fuel orifice 142 and a generally annular outerwall 144 that defines a secondary fuel orifice 146. The primary pilotsupply line 128 supplies fuel to the fuel nozzle 113 through the primaryfuel orifice 142, and the secondary pilot supply line 130 supplies fuelto the fuel nozzle 113 through the secondary fuel orifice 146.

As shown in FIGS. 4 and 5, the inner wall 140 is disposed radiallyinward with respect to the outer wall 144 such that the outer wall 144generally surrounds the inner wall 140 and the secondary fuel orifice146 surrounds the primary fuel orifice 142. Further, in the depictedembodiment, the primary fuel orifice 142 generally is radially alignedwith the secondary fuel orifice 146. That is, the primary and secondaryfuel orifices 142, 146 are disposed generally at the same axial locationwithin the fuel nozzle 113.

An annular pilot splitter 148 circumferentially surrounds the pilot fuelinjector 138. The pilot splitter 148 includes an upstream portion 150and a downstream portion 152. The upstream portion 150 generally iscylindrical in shape, while the downstream portion 152 generally isconical in shape. The downstream portion 152 generally is convergingwith respect to the centerline axis CL, having a wider first section 152a that gradually diminishes to a narrower second section 152 b, wherethe second section 152 b is downstream with respect to the first section152 a. A plurality of apertures 154 are defined in the second section152 b, e.g., the plurality of splitter apertures 154 may be definedalong the circumference of the second section 152 b and generally may beevenly spaced apart from one another. The splitter apertures 154 permita flow of air therethrough, e.g., to enhance cooling of the pilotsplitter 148 and thereby improve the splitter's durability. The flow ofair is described in greater detail below.

An annular outer boundary wall 156 circumferentially surrounds the pilotsplitter 148 and defines the outer boundary of a pilot portion P of thefuel nozzle 118. The outer boundary wall 156 includes a generallycylindrical first portion 156 a, a converging second portion 156 b, anda diverging third portion 156 c, such that a throat 158 is definedbetween the second and third portions 156 b, 156 c. As shown in FIG. 5,the first, second, and third portions 156 a, 156 b, 156 c are axiallyarranged in flow order, i.e., the first portion 156 a is upstream of thesecond portion 156 b, which is upstream of the third portion 156 c.Further, the converging second portion 156 b of the outer boundary wall156 generally follows or is parallel to the converging downstreamportion 152 of the pilot splitter 148. As such, a downstream end 160 ofthe pilot splitter 148 is disposed generally within the throat 158defined by the converging and diverging portions 156 b, 156 c of theouter boundary wall 156.

As illustrated in FIGS. 4 and 5, an inner air circuit 162 is definedbetween the pilot fuel injector 138 and the pilot splitter 148, and anouter air circuit 164 is defined between the pilot splitter 148 and theouter boundary wall 156. A circumferential array of inner swirl vanes166 radially extends from the pilot fuel injector 138 to the upstreamportion 150 of the pilot splitter 148. Similarly, a circumferentialarray of outer swirl vanes 168 radially extends from the upstreamportion 150 of the pilot splitter 148 to the first portion 156 a of theouter boundary wall 156. The inner swirl vanes 166 are shaped andoriented to induce a swirl into air flow passing through the inner aircircuit 162, and the outer swirl vanes 168 are shaped and oriented toinduce a swirl into air flow passing through the outer air circuit 164.

Upstream of the inner and outer air circuits 162, 164, the fuel nozzle113 defines a pilot air inlet 170 that permits ingress of air into thepilot portion P. The air flows into a pilot airflow passage 172, whichis split into the inner air circuit 162 and the outer air circuit 164 bythe pilot splitter 148. At the downstream end 160 of the pilot splitter148, the inner and outer air circuits 162, 164 merge back into thesingle pilot airflow passage 172, which extends through the remainder ofthe pilot portion P of the fuel nozzle 113. As shown in FIG. 4, thethird portion 156 c of the outer boundary wall 156 defines the outerboundary of the airflow passage 172 through the downstream end of thepilot portion P. The inner air circuit 162 and outer air circuit 164,including inner and outer swirl vanes 166, 168, and the third portion156 c of the outer boundary wall 156 form a pilot swirler 171. The pilotswirler 171 directs and controls the fluid flow, including the flow ofair and the mixture of air and fuel, through the pilot portion P of thefuel nozzle 113. More particularly, the air swirls through the inner andouter swirl vanes 166, 168 and then expands as it is mixed with fuel inthe generally conically shaped downstream portion of the pilot swirler171 defined by the outer boundary wall third portion 156 c.

Referring still to FIG. 4, a downstream end 174 of the outer boundarywall 156 may include a heat shield 176 that is configured as an annular,radially-extending plate. A thermal barrier coating (TBC) of a knowntype may be applied on all or a portion of the surface of the heatshield 176 and/or the outer boundary wall 156, e.g., to help protect thecomponents from the damaging effects of high temperatures. The heatshield 176 is described in greater detail below.

Further, the fuel nozzle 113 circumferentially surrounds the pilotportion P. In particular, an outer wall 121 of the fuel nozzle 113defines the fuel nozzle outlet 120 and extends axially to a radiallyoutermost end 178 of the heat shield 176. As illustrated in FIG. 4, theouter wall 121 is radially spaced apart from the outer boundary wall156. Additionally, the outer wall 121 defines an opening 123 thatpermits a flow of air into the space between the outer wall 121 and theouter boundary wall 156. The flow of air may provide cooling to the fuelnozzle heat shield 176.

The pilot fuel injector 138 defines a relatively small, stable pilotflame or burn zone. The pilot burn zone is centrally located within theannular combustor flow field in a radial sense. Fuel is supplied to thepilot fuel injector 138 via the primary and secondary pilot supply lines128, 130. Air is supplied through the pilot airflow passage 172. Thepilot airflow passage 172 provides a relatively high airflow; stateddifferently, the portion of the total combustor airflow directed throughthe pilot airflow passage 172 is relatively high, particularly comparedto known combustor designs. The airflow to and through the pilot portionP is described in greater detail below.

Continuing with FIG. 4, an annular main portion M extendscircumferentially about the annular pilot portion P of the fuel nozzle113. The main portion M includes a main fuel injector 180, which issupplied with fuel through a main fuel circuit 136. The main fuelcircuit 136 is coupled to and supplied with fuel by the main fuelconduit 134.

As illustrated in FIGS. 4, 6, 7, 8, and 9, the main fuel injector 180includes a plurality of injection ports 184. As generally provided inFIG. 4, the fuel nozzle 113 defines an angled main fuel injection port184 angled downstream with respect to the centerline axis CL of the fuelnozzle assembly 117. The fuel nozzle 113 further defines a radial mainfuel injection port 185 that is defined generally straight along theradial direction R relative to the centerline axis CL (e.g.,perpendicular to the centerline axis CL). Each of the angled fuelinjection ports 184 and radial fuel injection ports 185 are disposed inalternating arrangement through the main fuel injector 180. For example,the angled fuel injection port 184 may define every one port,alternating with each radial fuel injection port 185. As anotherexample, the angled fuel injection port 184 may define several ports inadjacent circumferential arrangement followed by several ports inadjacent circumferential arrangement of the radial fuel injection port185.

Each angled main injection port 184 of the fuel nozzle 113 has an inletend 186 and an outlet end 188, and the outlet end 188 is orienteddownstream with respect to the inlet end 186 and at an angle withrespect to the centerline axis CL. The inlet end 186 permits ingress offuel from the main fuel circuit 136 into the angled injection port 184,and the outlet end 188 permits egress of fuel from the injection port184. As such, the angled injection ports 184 of the fuel nozzle 113permit the egress of fuel from the main fuel circuit 136 toward thecenter of the combustion chamber 116 as described in greater detailbelow.

In the embodiment generally provided in FIG. 7, the fuel nozzle 113further defines a straight radial main fuel injection port 185 that isdefined straight along the radial direction R from the centerline axisCL (e.g., perpendicular to the centerline axis CL). That is, each radialinjection port 185 of the fuel nozzle 113 has an inlet end 186 and anoutlet end 188, and the outlet end 188 is oriented along the axialdirection A approximately equal with respect to the inlet end 186 andradially outward thereof with respect to the centerline axis CL.

The fuel nozzle assembly 117 further includes an annular main mixer orswirler 190 that circumferentially surrounds the fuel nozzle 113adjacent the main fuel injector 180. The main mixer 190 defines aplurality of inlet apertures 192 about its circumference to permitairflow into the main mixer 190. As shown in FIGS. 4, 6, 7, 8, and 9,the main mixer inlet apertures 192 are defined at a forward or upstreamend 194 of the main mixer 190. In some embodiments, the main mixer 190and its inlet apertures 192 may be shaped and/or oriented to induce aswirl into air flow passing through the main mixer 190. Downstream oraft of the apertures 192, the main mixer 190 includes an annular mainmixer wall 196 that extends to an aft or downstream end 198 of the mainmixer 190 and that is radially spaced apart from the outer wall 121 ofthe fuel nozzle 113. A main airflow passage 200 is defined between themain mixer wall 196 and the fuel nozzle outer wall 121. Further, themain mixer wall 196 defines a main mixer outlet 202 at the downstreamend 198. As such, air flows into the main mixer 190 through the inletapertures 192, continues through the main airflow passage 200, and exitsthe main mixer 190 through the main mixer outlet 202. The main mixer 190provides a relatively low airflow; stated differently, the portion ofthe total combustor airflow directed through the main mixer 190 isrelatively low, particularly compared to known combustor designs. Theairflow to and through the main portion M is described in greater detailbelow.

As also illustrated in FIGS. 4, 6, and 7, the fuel nozzle outer wall 121defines an aperture 204 therein that is aligned with the injection port184, 185. It will be appreciated that the outer wall 121 defines aplurality of apertures 204 that are each aligned with one of theinjection ports 184, 185. As previously stated, the angled injectionports 184 of the fuel nozzle 113 are angled downstream with respect tothe centerline axis CL of the fuel nozzle 113. The outer wall apertures204 similarly are defined at an angle with respect to the centerlineaxis CL; the angle of the apertures 204 may be substantially the same asthe angle of the angled injection ports 184 as shown in the exemplaryembodiment of FIGS. 4 and 6. Moreover, the outer wall apertures 204 aredefined downstream of the inlet apertures 192, such that the fuel isinjected within the main airflow passage 200 defined between the mainmixer wall 196 and the fuel nozzle outer wall 121. Accordingly, the fuelmixes in the main airflow passage 200 with the airflow introduced intothe main mixer 190 through the main mixer apertures 192, and thefuel-air mixture continues to flow downstream and exits the main mixer190 into the combustion chamber 116 through the main mixer outlet 202.As previously described, the angled injection ports 184 and outlet wallapertures 204 help direct the fuel toward the middle of the combustor101, such that the fuel within the combustor is more concentrated towarda center of the combustor. As such, the angled fuel injection may helpcontrol the profile and/or pattern factor of the combustor 101, as wellas allow a higher power operation of the engine and increase thedurability of the inner and outer liners 102, 104 and other combustorhardware by directing the fuel and combustion gases away from thecombustor hardware.

In other embodiments, the angled injection ports 184 may be angled in oralong other directions. For example, referring to FIG. 10, the angledinjection ports 184 are angled circumferentially around the fuel nozzle118, i.e., generally extending along the radial direction R but alsoalong the circumferential direction C as well as either upstream ordownstream along the axial direction A. As such, the angled injectionports 184 generally are aligned with the swirl direction of the mainmixer 190 or are perpendicular to the swirl direction of main mixer 190.As another example, illustrated in FIG. 11, the angled injection ports184 are angled upstream, rather than downstream as depicted in FIGS. 4,6, and 7. That is, the outlet end 188 of each angled injection port 184is oriented upstream with respect to the inlet end 186 and at an anglewith respect to the centerline axis CL. It will be appreciated that, asshown in FIGS. 6 and 7, the outer wall apertures 204 are defined toalign with the angled injection ports 184, no matter the orientation ofthe angled injection ports 184.

Further, it will be understood that the angled injection ports 184 havean orientation that is not purely or solely radial, axial, orcircumferential but, rather, comprises at least two directionalcomponents. In other words, because the ports 184 are angled, eachinjection port 184 does not extend along only the radial direction R,the axial direction A, or the circumferential direction C but extends,to some extent, along at least two directions. For example, referring toFIGS. 4 and 6, the orientation of angled injection ports 184 has aradial component as well as an axial component. That is, while eachangled injection port 184 of the depicted embodiment extends primarilyradially, the injection ports 184 are angled downstream such that theports 184 also extend in the downstream axial direction A. In theembodiment of FIG. 10, the fuel injection ports 184 extend in the radialdirection R, circumferential direction C, and axial direction A, and inthe embodiment of FIG. 11, the angled injection ports 184 extendradially as well as in the upstream axial direction A.

Referring still to FIG. 4, as well as the detailed view provided in FIG.7, a radial main injection port 185 of the fuel nozzle 113, includingthe inlet end 186 and the outlet end 188, is disposed in the radialdirection extended from the reference centerline axis CL. The fuelnozzle outer wall 121 defines the aperture 204 therein that is alignedwith the radial injection port 185 along the radial direction R. It willbe appreciated that the outer wall 121 defines a plurality of apertures204 that are each aligned with one of the radial injection ports 185. Aspreviously stated, the radial injection ports 185 of the fuel nozzle 113are disposed straight along the radial direction R (i.e., perpendicular)with respect to the centerline axis CL of the fuel nozzle 113. The outerwall apertures 204 similarly are defined straight with respect to thecenterline axis CL. Moreover, the outer wall apertures 204 are defineddownstream of the inlet apertures 192, such that the fuel is injectedwithin the main airflow passage 200 defined between the main mixer wall196 and the fuel nozzle outer wall 121. Accordingly, the fuel mixes inthe main airflow passage 200 with the airflow introduced into the mainmixer 190 through the main mixer apertures 192, and the fuel-air mixturecontinues to flow downstream and exits the main mixer 190 into thecombustion chamber 116 through the main mixer outlet 202. As previouslydescribed, the combination of the straight radial injection ports 185(FIG. 7) and the angled injection ports 184 (FIG. 6) and theirrespective outlet wall apertures 204 help provide non-uniform flamestructure and characteristics around the annulus of the combustionchamber 116 such as to de-couple the heat release and pressureoscillations, thereby mitigating combustion dynamics.

In another embodiment of the combustion system 100, the plurality offuel nozzles 113 defines a first fuel nozzle 118 and a second fuelnozzle 115 in circumferential arrangement in the combustion system 100.The first fuel nozzle 118, such as generally provided in FIG. 9, definesthe angled main fuel injection port 184 through the main fuel injection180 structure. The second fuel nozzle 115, such as generally provided inFIG. 8, defines the radial main fuel injection port 185 through the mainfuel injection 180 structure. The embodiments of the first fuel nozzle118 and the second fuel nozzle 115 generally provided may be configuredsubstantially similarly to the fuel nozzle 113 described in regard toFIGS. 4-7, and FIGS. 10-11. For example, identical numerals or referencenumbers indicated in regard to the fuel nozzle 113, and descriptions andembodiments thereof, are the same elements throughout the figures inregard to the first fuel nozzle 118 and the second fuel nozzle 115.

The first fuel nozzle 118 defining the angled main injection port 184(e.g., shown and described in regard to FIGS. 4, 6, and 9) and thesecond fuel nozzle 115 (e.g., shown in FIGS. 7 and 8) defining thestraight radial main injection port 185, in which the first fuel nozzle118 and the second fuel nozzle 115 are together in alternatingcircumferential arrangement around the longitudinal centerline 12 of theengine 10, produces a non-uniform flame structure and characteristicwhen fuel egresses from each main injection port 184 such as to mitigateundesired combustion dynamics. The non-uniform flame produced by thealternating arrangement of the first fuel nozzle 118 and the second fuelnozzle 115 de-couples the heat release oscillation from pressureoscillations, thereby mitigating undesired combustion dynamics (e.g.,vibrations, acoustics, noise, etc.).

The plurality of the fuel nozzle 113, as well as the combination of thefirst fuel nozzle 118 and the second fuel nozzle 115, with the fuelsystem 122 as described in regard to FIG. 2 may provide generallypassive combustion dynamics attenuation, in which the alternation of theangled main injection port 184 and the straight radial main injectionport 185 provides non-uniform flame structures and characteristic timesgiven an approximately equal flow of fuel to each main injection port184, 185 of each fuel nozzle 113. Furthermore, the angled main injectionport 184, the radial main injection port 185, the inlet end 186, theoutlet end 188, the outer wall apertures 204, or combinations thereof ofthe fuel nozzle 113 may define different volumes between one another. Asanother example, the angled main fuel injection port 184 of the firstfuel nozzle 118 may define a different volume or cross sectional area incontrast to the radial main fuel injection port 185 of the second fuelnozzle 115 such as to provide a different pressure and flow rate of fuelthrough the main injection port 184, 185 to the main airflow passage200.

In another embodiment, the plurality of fuel nozzles 113, as well as thecombination of the first fuel nozzle 118 and the second fuel nozzle 115,with the fuel system 122 as described in regard to FIG. 3 may providegenerally active combustion dynamics attenuation, in which thealternation of the angled main injection port 184 and the straightradial main injection port 185, in addition to variable flow, pressure,or temperatures of the fuel to each fuel nozzle 113, providesnon-uniform flame structures and characteristic times.

As previously described, the exemplary fuel nozzle 113 includes a heatshield 176 that is configured as an annular, radially-extending plate,as most clearly shown in FIG. 12. The heat shield area, which extendsbetween the pilot portion P and main portion M of the fuel nozzleassembly 117, is a stabilization zone for the combustion reaction. Thatis, hot combustion gases cross between the pilot portion P and the mainportion M to stabilize the reaction and keep the fuel burning properly.Thus, the hot gases are transported across the aft or outlet end 119 ofthe fuel nozzle 113 and the heat shield 176 helps to protect the outletend 119 of the fuel nozzle 113.

As depicted in FIGS. 4 and 6-9, the exemplary heat shield 176incorporates features for improving the durability of the heat shield,as it is exposed to the hot combustion gases. For instance, a radiallysealed cavity 206 is formed between the heat shield 176 and an aft end208 of the main fuel circuit 136. The cavity 206 receives a flow of airthrough apertures 210 defined in the aft end 208 of the main fuelcircuit 136. More particularly, airflow through the opening 123 definedby the fuel nozzle outer wall 121 may flow downstream within the spacebetween the fuel nozzle outer wall 121 and the outer boundary wall 156of the fuel nozzle pilot portion P. The airflow may continue through theapertures 210 and into the cavity 206 between the main fuel circuit 136and the heat shield 176. Further, the airflow into the cavity 206 mayimpinge on a forward surface 212 of the heat shield 176, which may helpcool the heat shield 176.

Moreover, as shown particularly in FIG. 12, the heat shield 176 definesone or more apertures 214 therein, through which the air may flow fromthe cavity 206 to an aft surface 216 of the heat shield 176. The heatshield apertures 214 may be angled, e.g., generally defined as passagesswirling into and out of the page in the schematic depictions of FIGS.4, 8, and 9, to lay a film of air along the aft surface 216 of the heatshield 176 and thereby help cool the aft surface 216. That is, coolingflow provided through heat shield apertures 214 may be swirled tocomplement the airflow local to the heat shield 176, which may create amore effective cooling film on the aft surface 216 of the heat shield176 without disrupting the flame stabilization zone. The combination ofimpingement and film cooling improves the durability of the heat shield176, which is exposed to hot combustion gases as described above.Additionally or alternatively, the heat shield apertures 214 may beshaped to reduce an exit velocity of the cooling flow, as well as tofurther improve film cooling of the heat shield 176. Further, a radialcompound angle may be employed to cool the radially outermost end 178 ofthe heat shield 176. The heat shield 176 also may incorporate otherfeatures for cooling the heat shield and improving its durability.

The fuel nozzle 113 and its constituent components, as well as the mainmixer 190, may be constructed from one or more metallic alloys.Non-limiting examples of suitable alloys include nickel and cobalt-basedalloys. All or part of the fuel nozzle 113 or portions thereof may bepart of a single unitary, one-piece, or monolithic component, and may bemanufactured using a manufacturing process that involves layer-by-layerconstruction or additive fabrication (as opposed to material removal aswith conventional machining processes). Such processes may be referredto as “rapid manufacturing processes” and/or “additive manufacturingprocesses,” with the term “additive manufacturing process” generallyreferring herein to such processes. Additive manufacturing processesinclude, but are not limited to: Direct Metal Laser Melting (DMLM);Laser Net Shape Manufacturing (LNSM); electron beam sintering; SelectiveLaser Sintering (SLS); 3D printing, such as by inkjets and laserjets;Stereolithography (SLA); Electron Beam Melting (EBM); Laser EngineeredNet Shaping (LENS); and Direct Metal Deposition (DMD). Other additive ornon-additive manufacturing processes may be used as well.

As previously stated, the pilot flow passage 172, or the pilot swirler171, provides a relatively high airflow while the main mixer 190provides a relatively low airflow. In some embodiments, the pilotswirler 171 provides an airflow of greater than about 14% W₃₆, where W₃₆is the total combustor airflow or total airflow into the combustorsystem 100. In particular embodiments, the pilot swirler 171 provides anairflow between about 15% W₃₆ to about 40% W₃₆, but the pilot swirler171 may provide a different amount of airflow as well. On the otherhand, the main mixer 190 provides an airflow of less than about 50% W₃₆.In particular embodiments, the main mixer 190 provides an airflowbetween about 25% W₃₆ to about 50% W₃₆, but the main mixer 190 mayprovide a different amount of airflow as well.

To provide a higher airflow, the size of the pilot air inlet 170 andpilot flow passage 172 are increased. For example, the pilot flowpassage 172 may have an increased radial height H_(p) with respect tothe fuel nozzle centerline axis CL. As such, the inner air circuit 162and/or outer air circuit 164 may have an increased radial height suchthat the inner and/or outer swirl vanes 166, 168 also have an increasedradial height. Generally, for a given operating condition of the engine10, a 100% increase in the area of the pilot flow passage 172 normal tothe air flowpath corresponds to a 100% increase in the percentage of thetotal combustor airflow to the pilot swirler 171. As an example, a knownpilot swirler design may have a pilot airflow at a high power operatingcondition of 10% W₃₆, with a flow passage area, normal to the directionof airflow, of X. Increasing the flow passage area, normal to thedirection of airflow, by 100% to 2X generally increases the pilotairflow at the high power operating condition to 20% W₃₆. Further, byutilizing CMC inner and outer liners 102, 104 to form the combustor 101of the combustion assembly 100, less cooling airflow is needed in thecombustor portion of the combustor system because CMC materials canwithstand higher temperatures than other typical combustor linermaterials, such as metallic materials. As such, less of the totalairflow to the combustor 101 is needed to cool the liners 102, 104, suchthat more of the total combustor airflow is available to the pilotswirler 171 and main mixer 190. Therefore, the additional availableairflow may be channeled through the pilot swirler 171 to increase theairflow through the pilot swirler, and the higher airflow through thepilot swirler 171 may be enabled by the pilot swirler design, e.g.,through an increased area of pilot flow passage 172.

Conversely, to reduce or lower the main mixer airflow, the size of themain airflow passage 200 is decreased. For instance, the main mixer wall196 is radially closer to the fuel nozzle outlet wall 121, whichdecreases the area of the flow passage 200 normal to the air flowpath bydecreasing the radial height of the flow passage 200. As described withrespect to increasing the area of the pilot flow passage 172, for agiven operating condition of the engine 10, a 50% decrease in the areaof the main flow passage 200 normal to the air flowpath generallycorresponds to a 50% decrease in the percentage of the total combustorairflow to the main mixer 190.

Increasing the airflow to the pilot swirler 171, particularly duringhigh power engine operations, may enable a different fuel split betweenthe pilot fuel injector 138 and the main fuel injector 180, compared toknown combustor system designs. In combustors, at least a portion of thefuel is distributed to the pilot fuel injector 138 at each engineoperating condition, i.e., the pilot portion P of the fuel nozzle 113 isconstantly supplied with fuel during engine operation. The portion ofthe fuel provided to the pilot fuel injector 138 may vary depending onthe engine operating condition. For example, at start up and low poweroperating conditions, 100% of the fuel may go to the pilot fuel injector138, while a lower percentage of the fuel goes to the pilot fuelinjector 138 and the remainder to the main fuel injector 180 at highpower conditions. Various transition fueling percentages may be used atpower levels in between low power and high power.

Known combustors provide a small fraction of the combustor airflow tothe pilot swirler, e.g., 10-13% W₃₆, such that the combustion systemwould not operate well at high power operating conditions if arelatively large portion of the fuel went to the pilot fuel injector.Typically, 10-20% of the fuel goes to the pilot fuel injector and 80-90%of the fuel goes to the main fuel injector at high power operatingconditions because the main mixer, with its higher airflow in a typicalcombustor, provides better fuel/air mixing and reduced NO_(x) emissions.However, a combustor incorporating the present subject matter asdescribed herein, namely, a high airflow pilot swirler 171, can providea much higher percentage of the fuel to the pilot fuel injector 138 athigh power operating conditions because of the higher pilot airflow. Thecombustor system 100 described herein may enable up to 100% of the fuelthrough the pilot fuel injector 138 over the full range of engineoperation. In some embodiments, the pilot fuel flow is within a range ofabout 30% to about 100% at high power, such that about 0% to about 70%of the fuel goes to the main injection ports 184 of the main fuelinjector 180. High pilot fuel flow may reduce combustion dynamics, i.e.,pressure oscillations in the combustor 101, and such high pilot fuelflows are made possible by the high pilot airflow split, where more airis available to mix with the fuel. As such, the combustor system 100described herein allows reduced combustion dynamics, improved fuel/airmixing, and reduced NO_(x) emissions. Further, as previously described,these and other features of the present combustor system 100 may helpimprove combustion efficiency, improve the durability of the fuel nozzle113 and combustor liners 102, 104, reduce smoke emissions, and improvethe profile/pattern factor of the engine.

As previously described, the inner liner 102 and outer liner 104 may beformed from a ceramic matrix composite (CMC) material, which is anon-metallic material having high temperature capability. In someembodiments, the combustor dome 114 also may be formed from a CMCmaterial. More particularly, the combustor dome 114 may be integrallyformed with the inner liner 102 and/or outer liner 104 from a CMCmaterial, such that the combustor dome 114 and the inner liner 102and/or outer liner 104 are a single piece. In other embodiments, thecombustor dome 114 may be formed separately from the inner and outerliners, either as a separate CMC component or from another suitablematerial, such as a metal or metal alloy. As described above, it may beparticularly useful to utilize CMC materials due to the relatively hightemperatures of the combustion gases 66, and the use of CMC materialswithin the combustor system 100 may allow reduced cooling airflow to theCMC components. However, other components of engine 10, such ascomponents of HP compressor 24, HP turbine 28, and/or LP turbine 30,also may comprise a CMC material.

Exemplary CMC materials utilized for such components may include siliconcarbide (SiC), silicon, silica, or alumina matrix materials andcombinations thereof. Ceramic fibers may be embedded within the matrix,such as oxidation stable reinforcing fibers including monofilaments likesapphire and silicon carbide (e.g., Textron's SCS-6), as well as rovingsand yarn including silicon carbide (e.g., Nippon Carbon's NICALON®, UbeIndustries' TYRANNO®, and Dow Corning's SYLRAMIC®), alumina silicates(e.g., Nextel's 440 and 480), and chopped whiskers and fibers (e.g.,Nextel's 440 and SAFFIL®), and optionally ceramic particles (e.g.,oxides of Si, Al, Zr, Y, and combinations thereof) and inorganic fillers(e.g., pyrophyllite, wollastonite, mica, talc, kyanite, andmontmorillonite). For example, in certain embodiments, bundles of thefibers, which may include a ceramic refractory material coating, areformed as a reinforced tape, such as a unidirectional reinforced tape. Aplurality of the tapes may be laid up together (e.g., as plies) to forma preform component. The bundles of fibers may be impregnated with aslurry composition prior to forming the preform or after formation ofthe preform. The preform may then undergo thermal processing, such as acure or burn-out to yield a high char residue in the preform, andsubsequent chemical processing, such as melt-infiltration or chemicalvapor infiltration with silicon, to arrive at a component formed of aCMC material having a desired chemical composition. In otherembodiments, the CMC material may be formed as, e.g., a carbon fibercloth rather than as a tape.

More specifically, examples of CMC materials, and particularlySiC/Si—SiC (fiber/matrix) continuous fiber-reinforced ceramic composite(CFCC) materials and processes, are described in U.S. Pat. Nos.5,015,540; 5,330,854; 5,336,350; 5,628,938; 6,024,898; 6,258,737;6,403,158; and 6,503,441, and U.S. Patent Application Publication No.2004/0067316. Such processes generally entail the fabrication of CMCsusing multiple pre-impregnated (prepreg) layers, e.g., the ply materialmay include prepreg material consisting of ceramic fibers, woven orbraided ceramic fiber cloth, or stacked ceramic fiber tows that has beenimpregnated with matrix material. In some embodiments, each prepreglayer is in the form of a “tape” comprising the desired ceramic fiberreinforcement material, one or more precursors of the CMC matrixmaterial, and organic resin binders. Prepreg tapes can be formed byimpregnating the reinforcement material with a slurry that contains theceramic precursor(s) and binders. Preferred materials for the precursorwill depend on the particular composition desired for the ceramic matrixof the CMC component, for example, SiC powder and/or one or morecarbon-containing materials if the desired matrix material is SiC.Notable carbon-containing materials include carbon black, phenolicresins, and furanic resins, including furfuryl alcohol (C₄H₃OCH₂OH).Other typical slurry ingredients include organic binders (for example,polyvinyl butyral (PVB)) that promote the flexibility of prepreg tapes,and solvents for the binders (for example, toluene and/or methylisobutyl ketone (MIBK)) that promote the fluidity of the slurry toenable impregnation of the fiber reinforcement material. The slurry mayfurther contain one or more particulate fillers intended to be presentin the ceramic matrix of the CMC component, for example, silicon and/orSiC powders in the case of a Si—SiC matrix. Chopped fibers or whiskersor other materials also may be embedded within the matrix as previouslydescribed. Other compositions and processes for producing compositearticles, and more specifically, other slurry and prepreg tapecompositions, may be used as well, such as, e.g., the processes andcompositions described in U.S. Patent Application Publication No.2013/0157037.

The resulting prepreg tape may be laid-up with other tapes, such that aCMC component formed from the tape comprises multiple laminae, eachlamina derived from an individual prepreg tape. Each lamina contains aceramic fiber reinforcement material encased in a ceramic matrix formed,wholly or in part, by conversion of a ceramic matrix precursor, e.g.,during firing and densification cycles as described more fully below. Insome embodiments, the reinforcement material is in the form ofunidirectional arrays of tows, each tow containing continuous fibers orfilaments. Alternatives to unidirectional arrays of tows may be used aswell. Further, suitable fiber diameters, tow diameters, andcenter-to-center tow spacing will depend on the particular application,the thicknesses of the particular lamina and the tape from which it wasformed, and other factors. As described above, other prepreg materialsor non-prepreg materials may be used as well.

After laying up the tapes or plies to form a layup, the layup isdebulked and, if appropriate, cured while subjected to elevatedpressures and temperatures to produce a preform. The preform is thenheated (fired) in a vacuum or inert atmosphere to decompose the binders,remove the solvents, and convert the precursor to the desired ceramicmatrix material. Due to decomposition of the binders, the result is aporous CMC body that may undergo densification, e.g., melt infiltration(MI), to fill the porosity and yield the CMC component. Specificprocessing techniques and parameters for the above process will dependon the particular composition of the materials. For example, silicon CMCcomponents may be formed from fibrous material that is infiltrated withmolten silicon, e.g., through a process typically referred to as theSilcomp process. Another technique of manufacturing CMC components isthe method known as the slurry cast melt infiltration (MI) process. Inone method of manufacturing using the slurry cast MI method, CMCs areproduced by initially providing plies of balanced two-dimensional (2D)woven cloth comprising silicon carbide (SiC)-containing fibers, havingtwo weave directions at substantially 90° angles to each other, withsubstantially the same number of fibers running in both directions ofthe weave. The term “silicon carbide-containing fiber” refers to a fiberhaving a composition that includes silicon carbide, and preferably issubstantially silicon carbide. For instance, the fiber may have asilicon carbide core surrounded with carbon, or in the reverse, thefiber may have a carbon core surrounded by or encapsulated with siliconcarbide.

Other techniques for forming CMC components include polymer infiltrationand pyrolysis (PIP) and oxide/oxide processes. In PIP processes, siliconcarbide fiber preforms are infiltrated with a preceramic polymer, suchas polysilazane and then heat treated to form a SiC matrix. Inoxide/oxide processing, aluminum or alumino-silicate fibers may bepre-impregnated and then laminated into a preselected geometry.Components may also be fabricated from a carbon fiber reinforced siliconcarbide matrix (C/SiC) CMC. The C/SiC processing includes a carbonfibrous preform laid up on a tool in the preselected geometry. Asutilized in the slurry cast method for SiC/SiC, the tool is made up ofgraphite material. The fibrous preform is supported by the toolingduring a chemical vapor infiltration process at about 1200° C., wherebythe C/SiC CMC component is formed. In still other embodiments, 2D, 2.5D,and/or 3D preforms may be utilized in MI, CVI, PIP, or other processes.For example, cut layers of 2D woven fabrics may be stacked inalternating weave directions as described above, or filaments may bewound or braided and combined with 3D weaving, stitching, or needling toform 2.5D or 3D preforms having multiaxial fiber architectures. Otherways of forming 2.5D or 3D preforms, e.g., using other weaving orbraiding methods or utilizing 2D fabrics, may be used as well.

Thus, a variety of processes may be used to form a CMC inner liner 102and a CMC outer liner 104, as well as any other CMC components of thecombustor system 100, such as combustor dome 114, and/or engine 10. Ofcourse, other suitable processes, including variations and/orcombinations of any of the processes described above, also may be usedto form CMC components for use with the various combustor systemembodiments described herein.

As described herein, the present subject matter provides combustorsystems having different airflow and fuel splits than known combustorsystems. In particular, the present subject matter provides a relativelyhigher pilot swirler airflow and a relatively lower main mixer airflow,which allows a higher fuel flow to the pilot portion P of the fuelnozzle 113, particularly during high engine power operations. Thedifferent airflow splits may be enabled through the use of CMC combustorliners 102, 104, which require less cooling airflow than combustorliners made from different materials, such as metallic materials. Thepresent subject matter also provides downstream angled fuel injectionthrough the main fuel injector 180, which may help improve thedurability of the downstream combustor components, such as the combustorliners 102, 104, as well as allow higher power operation of the engine.Further, in some embodiments, the angled fuel injection ports 184 may beformed by additively manufacturing the main fuel circuit 136, whichmanufacturing process may help precisely define the fuel injection ports184. Moreover, the present subject matter provides cooling or purgeholes through the pilot splitter 148, which may help improve thedurability of the pilot splitter. As such, the combustor systems andfuel nozzle assemblies described herein allow engine operation at arelatively high fuel/air stoichiometry with high combustion efficiency,reduced or low combustion dynamics, improved fuel nozzle and combustorliner durability, low smoke and NO_(x) emissions, and a reduced or lowprofile and pattern factor. The present subject matter may have otherbenefits and advantages as well.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal language of the claims.

What is claimed is:
 1. A fuel nozzle assembly comprising: a fuel nozzlecomprising a main fuel injector defining a centerline axis, a main mixeraround the centerline axis defining an annular main mixer wall, a mainairflow passage defined between the main fuel injector and the annularmain mixer wall, the main fuel injector defining an angled maininjection port, the angled main injection port defining an angled maininjection port inlet end and an angled main injection port outlet end,wherein the angled main injection port outlet end is positioneddownstream of the angled main injection poll inlet end relative to afuel flow through the angled main injection port, wherein the angledmain injection port is positioned at an angle relative to the centerlineaxis, wherein the angled main injection port is further positioned at anangle in a circumferential direction around the centerline axis suchthat the angled main injection port is non-orthogonal to the centerlineaxis, wherein the angled main injection port permits an egress of fuelfrom a main fuel circuit into the main airflow passage, and the mainfuel injector further defining a radial main injection port, the radialmain injection port defining a radial main injection port inlet end anda radial main injection port outlet end, wherein the radial maininjection port is orthogonal to the centerline, wherein the radial maininjection port permits an egress of fuel into the main airflow passage,and wherein the angled main injection port outlet end is positionedupstream or downstream of the radial main injection port outlet endrelative to an airflow through the main airflow passage of the mainmixer.
 2. The fuel nozzle assembly of claim 1, wherein the fuel nozzledefines an annular fuel nozzle outer wall radially spaced apart from anannular outer boundary wall surrounding a pilot fuel injector, whereinthe annular fuel nozzle outer wall defines an opening that permits aflow of air between the annular fuel nozzle outer wall and the annularouter boundary wall.
 3. The fuel nozzle assembly of claim 1, wherein thefuel nozzle defines an aperture through the annular fuel nozzle outerwall, wherein the aperture is defined in alignment with the angled maininjection port, and wherein the aperture defines an angle substantiallyequal to the angle of the angled main injection port.
 4. The fuel nozzleassembly of claim 2, wherein the fuel nozzle defines an aperture throughthe annular fuel nozzle outer wall, wherein the aperture is defined inalignment with the radial Main injection port.
 5. The fuel nozzleassembly of claim 1, wherein the fuel nozzle comprises a pilot fuelinjector, wherein the pilot fuel injector comprises an pilot fuelinjector annular inner wall defining a primary fuel orifice and a pilotfuel injector annular outer wall defining a secondary fuel orifice, andwherein a primary pilot supply line supplies fuel through the primaryfuel orifice and a secondary pilot supply line supplies fuel through thesecondary fuel orifice.
 6. The fuel nozzle assembly of claim 5, furthercomprising: a fuel system comprising a pilot control valve coupled to apilot fuel conduit and operable to supply a flow of fuel thereto,wherein the pilot fuel conduit is in fluid communication with theprimary pilot supply line and the secondary pilot supply line.
 7. Thefuel nozzle assembly of claim 6, wherein the fuel system furthercomprises a first main valve fluidly coupled to a first main fuelconduit and operable to supply a flow of fuel thereto, wherein the firstmain fuel conduit is in fluid communication with a first the main fuelcircuit of the fuel nozzle.
 8. The fuel nozzle assembly of claim 7,wherein the fuel nozzle defines a different volume of the angled mainfuel injection port relative to the radial main fuel injection port, orwherein the fuel nozzle defined a different cross sectional area of theangled main fuel injection port inlet end, or the angled main fuelinjection port outlet end, relative to the radial main fuel injectionport inlet end, or the radial main fuel injection port outlet end. 9.The fuel nozzle assembly of claim 1, Wherein the fuel nozzle furthercomprises a pilot fuel injector, wherein each of the main fuel injectorand the pilot fuel injector are configured to receive a portion of afuel flow.
 10. A gas turbine engine, comprising: a combustor systemcomprising a combustion chamber; and a plurality of fuel nozzles, theplurality of fuel nozzles comprising: a first fuel nozzle comprising: amain fuel injector defining a centerline axis, a main mixer around thecenterline axis defining an annular main mixer wall, a main airflowpassage defined between the main fuel injector and the annular mainmixer wall, the main fuel injector defining an angled main injectionport, the angled main injection port defining an angled main injectionport inlet end and an angled main injection port end, wherein the angledmain injection port outlet end is positioned downstream of the angledmain injection port inlet end relative to a fuel flow through the angledmain injection port, wherein the angled main injection port ispositioned at an angle relative to the centerline axis, wherein theangled main injection port is further positioned at an angle in acircumferential direction around the centerline axis such that theangled main injection port is non-orthogonal to the centerline axis,wherein the angled main injection port permits an egress of fuel from amain fuel circuit into the main airflow passage, and the main fuelinjector further defining a radial main injection port the radial mainport outlet end, wherein the radial main injection port is orthogonal tothe centerline, wherein the radial main injection port permits an egressof fuel into the airflow passage, and wherein the angled main injectionport outlet end is positioned upstream or downstream of the radial maininjection port outlet end relative to an airflow through the mainairflow passage of the main mixer; and a second fuel nozzle comprising asecond radial main injection port defining a second radial maininjection port inlet end and a second radial main injection port outletend, and wherein the second radial main injection port outlet end andthe second radial main injection port inlet end are located in a planeperpendicular to a second centerline axis.
 11. The gas turbine engine ofclaim 10, further comprising: a fuel system comprising a pilot controlvalve coupled to a pilot fuel conduit and operable to supply a flow offuel thereto, wherein the pilot fuel conduit is in fluid communicationwith a primary pilot supply line and a secondary pilot supply line. 12.The gas turbine engine of claim 11, wherein the fuel system furthercomprises a first main valve coupled to a first main fuel conduit andoperable to supply a flow of fuel thereto, wherein the first main fuelconduit is in fluid communication with the first main fuel circuit. 13.The combustor system of claim 12, wherein the first fuel nozzle definesa different cross sectional area of the angled main fuel injection port,the angled main injection port inlet end, or the angled main injectionport outlet end, relative to a cross sectional area of the second radialmain fuel injection port, the second radial main fuel injection portinlet end, or the second radial main fuel injection port outlet end. 14.The gas turbine engine of claim 10, wherein the fuel system furthercomprises: a first main valve fluidly coupled to a first main fuelconduit and operable to supply a flow of fuel thereto, wherein the firstmain fuel conduit is in fluid communication with the first main fuelcircuit; and a second main valve fluidly coupled to a second main fuelconduit and operable to supply a flow of fuel thereto independent of thefirst main fuel conduit, wherein the second main fuel conduit is influid communication with a second main fuel circuit at the second fuelnozzle.